This invention relates generally to gas turbine engines, and more particularly, to methods and apparatus for operating gas turbine engines.
At least some known gas turbine engines typically include an inlet, a fan assembly, low and high pressure compressors, a combustor, and at least one turbine. The compressors compress air which is channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) which extracts energy from the combustion gases for powering the compressor(s), as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
Some known fan assemblies include a casing that encloses a plurality of blades coupled to a fan rotor wherein such blades may be subject to events that facilitate at least partial fan blade breakage. Such breakage facilitates primary damage which includes the affected blade and the immediately downstream blades as they contact the material released from the affected blade. Such primary damage may induce rotor unbalancing conditions and subsequent blade rubs against the fan casing. The blade rubs may facilitate secondary damage that includes damage to non-adjacent blades and the casing.
Many known fan assemblies are designed with a sufficient margin of error and constructed with sufficient additional load-carrying capabilities to compensate for such unbalanced rotor conditions and reduce a potential for damage in blade breakage events. Such additional load-carrying capabilities increase a cost of construction of the fan assemblies and decrease a gas turbine engine fuel efficiency due to the increased weight of the fan assemblies.